This invention relates to the nosecone design of a rocket or aerospace plane. More specifically, this invention relates to the mitigation of the shock front experienced at hypersonic speed through supercooling.
U.S. Pat. No. 5,191,761, owned by the applicant for the present invention, discloses an air breathing aerospace engine. That patent is incorporated by reference in its entirety. The engine includes a frontal core that houses an oxygen liquefaction system that captures ambient air, liquefies and separates the oxygen. The oxygen may then be used in the rocket engine.
U.S. Pat. No. 6,213,431, owned by the applicant of the present invention, discloses an aerospike engine. That patent is incorporated by reference in its entirety. An aerospike engine may have a tapered body with a slanted or curved reaction plane. A fuel injector directs fuel down the reaction plane. The combustion of the fuel on the reaction plane creates a propulsive force across the reaction plane.
U.S. Pat. No. 7,344,111 hence discloses a re-usable or reversible SSTO that may be expediently launched to service the rapidly expanding space enterprise. Merging patent '761 and patent '431 rendered a unique aerospace plane confirmation (the '111 patent) with a liquefaction or supercooling nosecone on the ascending or leading direction and an adaptive aerospike rocket engine on the tail end.
Oxygen constitutes 89% of the mass of the necessary propellant for such an engine. A substantial portion of the necessary liquid oxygen can be distilled out of the ambient atmosphere at hypersonic speeds by supercooling through turbo compression at the nosecone. Through this process, the merged aerospace plane would be capable of tanking the necessary hydrogen as well as carrying a substantially improved payload into orbit. However, a consequence of tanking all of the necessary liquid hydrogen propellant onboard would be to invariably result in bubbly contours.
A reversible (“Uturn”) aerospace plane was rendered that can fly hypersonically without the impediment of shock waves, shock front, or superheated intake air (a node of optimality) in the ascending direction whilst reentering in the reversed (high drag) direction into the atmosphere with the benefit of cushioned heat dissipation in the higher/rarified atmospheric domain. The slated “turbocharged” aerospace plane may also be barrel-rolled through the initial reentry phase so as to dissipate the insipient heat into deep space. As a consequence of employing the cryogenic potential of the liquid oxygen/hydrogen propellants to supercool the nosecone in lieu of cooling the jacket of the rocket engine, it would be conversely be necessary to line the expansion ramp of the aerospike engine with ceramic tiles. Slated ceramic tiled would function both as a reentry shield as well as active ramp insulation. The Uturn would hence facilitate an aerospace plane with substantially lesser heat shield whilst limiting ceramic tiles to the aero spike engine expansion ramps.
Whereas the efficacy of the oxygen liquefaction would be substantially constrained by the extremely low pressure of the rarified atmospheric air at high altitudes with the “111” patent in its native form, the addition of a turbo compressor driven by either superheated nitrogen or superheated hydrogen in accordance with claims #1 and #13 of the “111” patent would substantially expand the operating envelope of the air breathing aerospace engine. The high compressed ambient air may hence be inter-cooled in accordance with the means provided via patent “111” and hence expanded/flashed into the cryogenic zone via an expansion motor/turbine that may either augment the compression process or drive the propellant pumps. Means of supercooling the nose cone at supersonic speeds is additionally being introduced to suppress the formation of the shock front in the abstract (eg lim dQ>>infinity, dL=0), “freezing” OR usurping the incipient shock at Mach 1 and rendering isothermal compression in lieu of adiabatic compression of the intake air into the liquefaction plant throughout the hypersonic regime.
A novel new approach with a rocket or aerospace plane flying hypersonically through the atmosphere entails morphing (cool) isothermal compression (in lieu of a blazing shock front) by orchestrating supercooling and superconductivity in coherence, usurping the insipient shock wave at formation (Prandt1 singularity, Mach1) and distilling both liquid oxygen as well as liquid nitrogen simultaneously out of the incident atmospheric air via regenerative chilling. The rational simply entails that in lieu of dissipating 99.99% of the incident kinetic energy in the shock front at hypersonic velocity (an irreversible adiabatic process), 99.99% of the kinetic energy is conversely converted into useful work via isothermal compression of the incident atmospheric air. By employing the cryogenic power of tanked hydrogen incident to a H2/O2 propulsive system coherently and morphing isothermal compression by means of the Prandt1 singularity “niche” via the force of supercooling AND superconductivity, compressed supercool/saturated/liquefacted ambient air may be rendered as a unique solution to a perplexing technological challenge.
The following definitions apply to terms used in this application:                Liquefaction: The condensation of a gaseous medium.        Supercooling: Rate of cooling orders of magnitude in excess of normal heat transfer rates.        Superconduction: Rate of conduction orders of magnitude in excess of normal conduction rates.        Superemissivity/absorptivity: Emissivity or absorptivity in excess of unity by means of negative refraction (or contact area morphing).        Supersonic: Flying 2-3 times the speed of sound.        Hypersonic: Flying 4-20 times the speed of sound.        Isothermal compression: Compression at constant temperature        Adiabatic compression: Compression without loss of heat (eg rapid compression via a shock wave).        
Normally compression of air results in an increase in the temperature of the incident air. Air may be compressed either via a mechanical device (reciprocating or turbo compressor) OR via the ram force of a body traveling though the atmosphere. Compression of the air may be enhanced via a diffuser in the latter instance. At high speeds rapid compression of the air results in formation of a shock wave (adiabatic/trapped heat of compression). The speed of the body traveling through the atmosphere at formation of a shock wave is denoted Mach1. The energy content of the air at Mach1 is labeled “total” or “stagnation” temperature “Ts” in absolute terms. The pertinent relationships are as follows:Ts=To[1+(k−1)/2×M^2]  (1)where Ts is the stagnation temperature, To=ambient temperature in absolute terms, k=polytropic constant Cp/Cv=1.4 and M=Mach number;T2/T1=(p2/p1)^k−1/k  (2)Where T2/T1 is the adiabatic temperature ratio and p2/p1 is the compression ratio;wa=Cp·R·T1·(k/k−1)[(p2/p1)^k−1/k−1]  (3)where wa is the adiabatic work of compression, R is the ideal gas constant, T entering temperature;wi=Cp·R·T1·ln(p2/p1)  (4)where wi is the isothermal work of compression with a compression ratio of p2/p1 vis-à-vis; Atmospheric pressure model (http://en.wikipedia.org/wiki/Atmospheric_pressure);
101/218,000ft1/327,480ft1/1052,926ft1/100101,381ft1/1,000159,013ft1/10,000227,899ft1/100,000283,076ft.                0/E^3=20; 50,000 ft/E^6=403; 95,000 ft/E^9=8,103; 150,000 ft/E^12=162,755The real time Mach number/stagnation temperature relationship in terms of altitude has been determined as follows (http://www.grc.nasa.gov/WWW/BGH/stagtmp.html);        
M1 = 500 R@5000ftM2 = 700 R@30-90,000ftM3 = 1100 R@40-90,000ftM4 = 1600 R@40-90,000ftM5 = 2200 R@40-100,000ftM6 = 3200 R@40-100,000ftM7 = 4200 R@40-100,000ftM8 = 5300 R@40-100,000ft.The respective shock/gradient for a (real time) adiabatic shock front (plus associated frontal pressure) may hence be rendered via the polytropic relationship p2/p1=[T2/T1]^k/(k−1) as follows;
M1: p2/p1 = 2.2p2 = paxp2/p1 = 2 × 15/2.75 = 11 psi@20,000ftM2: p2/p1 = 3.2p2 = paxp2/p1 = 3.2 × 15/4 = 12 psi@30,000ftM3: p2/p1 = 16M4: p2/p1 = 59M5: p2/p1 = 179p2 = paxp2/p1 = 179 × 15/55 = 49@70,000ftpsiM6: p2/p1 = 663M7: p2/p1 = 1718M8: p2/p1 = 3878×15/1100 = 53 psi@100,000ftIn event of isothermal compression the Bernoulli's law will control with p=V^2/2vg/144 in psi. The respective frontal pressure in event of isothermal compression hence becomes:
M1: p1 =M2: p2 = 10.3 psi@27,480ft (1/2)M3: p3 =M4: p4 = 3.9 psi@52,926ft (1/3)M5: p5 =M6: p6 = 2.8 psi@101,381ft (1/100)M7: p7 =M8: p8 = 2.0 psi@120,000ft (1/245)M9: p9 =M10: p10 = 0.77 psi@159,013ft (1/1000)The work of compression for a (normalized) free-range adiabatic compression hypersonic shock ram versus isothermal compression (in the abstract) at Mach8 is as follows:wa=Cp·R·T1·(k/k−1)[(p2/p1)^k−1/k−1]=0.25×3.5×53.3×400/788×[(3878)^0.286−1]=226 Btu/lbwi=Cp·R·T1·ln(p2/p1)=0.25×53.3×400/788×ln(245)=37Btu/lb.
The impact of a sustainable isothermal (ram) compression system thru the hypersonic domain is threefold, eg 1) that the destructive thermal impact in the native/adiabatic mode is being contained 2) the work of compression is substantially mitigated to the extent that the latent heat of evaporation of hydrogen component alone of the LH2/LO2 propellant will suffice in dissipating the incident heat of compression within the realm of the supercool synthesis and 3) the drag force of a “superconductive hypersonic liquefaction nosecone” will be proximal 6× (eg 226/37) less in comparison to a native hypersonic aerospace plane flying trough the hypersonic zone.
The mass flow rate at Mach8 (8,000 ft/sec) conversely=8,000×1/14/245=2.33/lb/SF/sec. The heat of compression hence becomes:
Qa=226×2.33×3600=1,895,688 Btu/SF/h
Qi=37×2.33×3600=310,356 Btu/SF/h.
O2: Latent Heat of Vaporization: 2934 BTU/lb mole (92 Btu/lb)
http://www-safety.deas.harvard.edu/services/oxygen.html#physical
N2: Latent Heat of Vaporization: 2399 BTU/lb mole (86 Btu/lb)
http://www-safety.deas.harvard.edu/services/nitrogen.html#physical
H2: Latent Heat of Vaporization: 389 Btu/lb-Mole (195 Btu/lb)
http://www.ehs.ufl.edu/Lab/Cryogens/hydrogen.html
H2 has two different nuclear spin states: ortho-hydrogen (spin 1 state) and para-hydrogen (spin 0 state). Near room temperature, equilibrium hydrogen is 75% ortho and 25% para. However, at low temperatures, around the normal boiling point of 20.3 K, hydrogen is nearly all in the para-hydrogen state. The conversion process from ortho to para hydrogen is exothermic and generates about 700 KJoule/mole.